1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine stator vane with sequential impingement cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, air is first compressed to a high pressure in a compressor. The high pressure air is then mixed with fuel and burned at nearly constant pressure in the combustor. The high temperature gas exhausted from the combustor is then expanded through a turbine which then drives the compressor. If executed correctly, the exhaust stream from the turbine maintains sufficient energy to provide useful work by forming a jet, such as in aircraft jet propulsion or through expansion in another turbine which may then be used to drive a generator like those used in electrical power generation. The efficiency and power output from these machines will depend on many factors including the size, pressure and temperature levels achieved and an agglomeration of the efficiency levels achieved by each of the individual components.
Current turbine components are cooled by circulating relatively (to the gas turbine engine) cool air, which is extracted from the compressor, within passages located inside the component to provide a convective cooling effect. In many recent arrangements, the spent cooling flow is discharged onto the surfaces of the component to provide an additional film cooling effect.
The challenge to cool first stage turbine vanes (these are exposed to the highest temperature gas flow), in particular, is complicated by the fact that the pressure differential between the vane cooling air and the hot gas which flows around the airfoil must necessarily be small to achieve high efficiency. Specifically, coolant for the first stage turbine vane is derived from the compressor discharge, while the hot gas is derived from the combustor exit flow stream. The pressure differential available for cooling is then defined by the extremely small pressure drop which occurs in the combustor. This is because the pressure of the coolant supplied to the vane is only marginally higher than the pressure of the hot gas flowing around the airfoil as defined by the combustor pressure loss, which is desirably small. This pressure drop is commonly on the order of only a few percentage points. Further, it is desirable to maintain coolant pressure inside the vane higher than the pressure in the hot gas flow path to insure coolant will always flow out of the vane and not hot gas into the vane. Conversely, in the event hot gas is permitted to flow into the vane, serious material damage can result as the materials are heated beyond their capabilities and progression to failure will be swift. As a consequence, current first stage turbine vanes are typically cooled using a combination of internal convection heat transfer using single impingement at very low pressure ratio, while spent coolant is ejected onto the airfoil surface to provide film cooling.
The efficiency of the convective cooling system is measured by the amount of coolant heat-up divided by the theoretical heat-up possible. In the limits, little coolant heat-up reflects low cooling efficiency while heating the coolant to the temperature of the surface to be cooled (a theoretical maximum) yields 100% cooling efficiency. In the previous methods using single impingement, the flow could only be used once to impinge on the surface to be cooled. This restriction precludes the ability to heat the coolant substantially, thereby limiting the cooling efficiency.
U.S. Pat. No. 8,096,766 issued to Downs on Jan. 17, 2012 and entitled AIR COOLED TURBINE AIRFOIL WITH SEQUENTIAL COOLING discloses one such insert having sequential cooling where the insert is build up from a stack of alternating plates that are bonded together in chordwise planes. This insert provides for the sequential cooling of the vane walls but is a very expensive insert because of the numerous plates that must be individually machined and then bonded together.